High temperature turbine nozzle for temperature reduction by optical reflection and process for manufacturing

ABSTRACT

A high temperature gas turbine component for use in the gas flow path that also is a specular optical reflector. A thin layer of a high temperature reflector is applied to the gas flow path of the component, that is, the surface of the component that forms a boundary for hot combustion gases. The component typically includes a thermal barrier coating overlying the high temperature metallic component that permits the component to operate at elevated temperatures. The thermal barrier coating must be polished in order to provide a surface that can suitably reflect the radiation into the gas flow path. A thin layer of the high temperature reflector the is applied over the polished thermal barrier coating by a process that can adequately adhere the reflector to the polished surface without increasing the roughness of the surface. The high temperature reflector can be applied to any surface aft of the compressor, such as on a turbine nozzle. The surface reflects radiation back into the hot gas flow path. The reflected radiation is not focused onto any other hardware component. The design of the component is such that the radiation is returned to the gas flow path rather than absorbed into a component wall that only serves to increase the temperature of the wall.

CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] This Application is related to application Ser. No.__/______Attorney Docket No. 13DV-13956, filed contemporaneously withthis Application on Dec. 31, 2002, entitled “IMPROVED HIGH TEMPERATURESPLASH PLATE FOR TEMPERATURE REDUCTION BY OPTICAL REFLECTION AND PROCESSFOR MANUFACTURING” assigned to the assignee of the present invention andwhich is incorporated herein by reference, to application Ser. No.__/______ Attorney Docket No. 13DV-13957 filed contemporaneously withthis Application on Dec. 31, 2002, entitled “IMPROVED HIGH TEMPERATURECENTERBODY FOR TEMPERATURE REDUCTION BY OPTICAL REFLECTION AND PROCESSFOR MANUFACTURING” assigned to the assignee of the present invention andwhich is incorporated herein by reference, and to application Ser. No.__/______ Attorney Docket No. 13DV-13959 filed contemporaneously withthis Application on Dec. 31, 2002, entitled “IMPROVED HIGH TEMPERATURECOMBUSTOR WALL FOR TEMPERATURE REDUCTION BY OPTICAL REFLECTION ANDPROCESS FOR MANUFACTURING” assigned to the assignee of the presentinvention and which is incorporated herein by reference.

FIELD OF THE INVENTION

[0002] The present invention is directed to gas turbine engines, and inparticular, to modifications of components of such engines to reduce thetemperature of boundary walls of the hot section portions of thecomponents by optical radiation generated by combustion

BACKGROUND OF THE INVENTION

[0003] In the compressor portion of an aircraft gas turbine engine,atmospheric air is compressed to 10-25 times atmospheric pressure, andadiabatically heated to about 800°-1250° F. (425°-675° C.) in theprocess. This heated and compressed air is directed into a combustor,where it is mixed with fuel. The fuel is ignited, and the combustionprocess heats the gases to very high temperatures, in excess of 3000° F.(1650° C.). These hot gases pass through the turbine, where rotatingturbine wheels extract energy to drive the fan and compressor of theengine, and the exhaust system, where the gases supply thrust to propelthe aircraft. To improve the efficiency of operation of the aircraftengine, combustion temperatures have been raised. Of course, as thecombustion temperature is raised, steps must be taken to prevent thermaldegradation of the materials forming the flow path for these hot gasesof combustion.

[0004] Every aircraft gas turbine engine has a so-called High PressureTurbine (HPT) to drive its compressor. The HPT sits just behind thecompressor in the engine layout and experiences the highest temperatureand pressure levels (nominally 2400° F. and 300 psia respectively)developed in the engine. The HPT also operates at very high speeds(10,000 RPM for large turbofans, 50,000 for small helicopter engines).In order to meet life requirements at these levels of temperature andpressure, today's HPT components are always air-cooled and constructedfrom advanced alloys.

[0005] While a straight turbojet engine will usually have only oneturbine (an HPT), most engines today are of the turbofan, either highbypass turbofan or low bypass turbofan, or turboprop type and requireone (and sometimes two) additional turbine(s) to drive a fan or agearbox. The additional turbines are called the Low Pressure Turbines(LPT) and immediately follows the HPT in the engine layout. Sincesubstantial pressure drop occurs across the HPT, the LPT operates with amuch less energetic fluid and will usually require several stages(usually up to six) to extract the available power.

[0006] One well-known solution that has been undertaken to protect themetals that form the flow path for the hot gases of combustion,including those of the HPT and LPT, have included application ofprotective layers having low thermal conductivity. These materials areapplied as thermal barrier coating systems (TBCs), typically comprisinga bond coat that improves adhesion of an overlying ceramic top coat,typically a stabilized zirconia, to the substrate. These systems areknown to improve the thermal performance of the underlying metals thatform the flow path in the hot section of the engine. However, astemperatures of combustion have increased, even these TBCs have beenfound to be insufficient.

[0007] Another solution that has been used in conjunction with TBCs isactive cooling of metal parts. Initially, active cooling provided a flowof air from the compressor to the back side of metal parts comprisingthe flow gas path. As temperatures increased even further, serpentinepassageways were formed in the metallic components and cooling air wascirculated through the parts to provide additional cooling capability,the cooling air exiting through apertures positioned in the gas flowside of the component, providing an additional impingement film layeralong the gas flow path. This method is referred to as film cooling.Even though the air from the compressor is adiabatically heated toperhaps as high as 1250° F. (675° C.), the compressor air is stillsignificantly cooler than the combustion gases moving along the gas flowpath of the engine. However, as the temperatures of the combustionprocess have continued to increase, even these tried and true methods ofcooling are reaching their limitations. In particular, the turbinenozzles of high efficiency, advanced cycle turbine engines are prone tofailure as a result of thermal degradation.

[0008] While some modifications of the traditional flow path surfaceshave been applied in the past, such as the application of materials overthe TBC, these modifications have been directed to reducing theemissions of pollutants such as unburned hydrocarbons (UHC) and carbonmonoxide (CO). One such modification is set forth in U.S. Pat. No.5,355,668 to Weil, et al., assigned to the assignee of the presentinvention, which teaches the application of a catalyst such as platinum,nickel oxide, chromium oxide or cobalt oxide directly over the flow pathsurface of the thermal barrier coating of a component such as a turbinenozzle. The catalyst layer is applied to selected portions of flow pathsurfaces to catalyze combustion of fuel. The catalytic material ischosen to reduce air pollutants such as unburned hydrocarbons (UHC) andcarbon monoxide (CO) resulting from the combustion process. Thecatalytic layer is applied to a thickness of 0.001 to 0.010 inches andis somewhat rough and porous, having a surface roughness of about 100 to250 micro inches, in order to enhance the surface area available tomaximize contact with the hot gases in order to promote the catalyticreaction. The rough surface assists in creating some turbulence thatpromotes contact with the catalytic surface.

[0009] The prior art solutions are either directed to problems that areunrelated to the problem of high temperature experienced by turbinenozzles, such as the Weil patent, or provide different solutions to theproblem of high temperatures resulting from the combustion process. Thepresent invention provides a different approach to the problem of hightemperatures experienced by turbine nozzles.

SUMMARY OF THE INVENTION

[0010] The present invention is a high temperature gas turbine componentfor use in the gas flow path that also is a specular optical reflector.The gas turbine component is positioned in the hot section of theengine, behind the compressor section and reflects heat radiation, forexample infrared radiation having wavelengths in the range of about 1micron to about 10 microns, from the combustor region back into the hotgas flow path. The reflected radiation is focused away from any otherhardware component in the combustor region so that the radiative heatpasses into the turbine portion of the engine. The design of thecomponent is such that the radiation is returned to the gas flow pathrather than absorbed into a component wall which only serves to increasethe temperature of the wall.

[0011] A thin layer of a high temperature reflector metal is applied tothe flow path surface of the component, that is, the surface of thecomponent that forms a boundary for hot combustion gases. The hightemperature reflector must be applied as an optically smooth coating.The component typically includes a thermal barrier coating overlying thehigh temperature metallic component that permits the component tooperate at elevated temperatures. The thermal barrier coating (TBC)applied to the component typically is rough and must be polished inorder to provide a sufficiently smooth surface that can suitably reflectthe radiation into the gas flow path. A thin layer of the hightemperature reflector then is applied by a process that can adequatelyadhere the reflector to the polished TBC surface without increasing theroughness of the surface. The high temperature reflector can be appliedto any surface aft of the compressor, but is most beneficially used inthe combustor portion of the engine, for instance, the combustor wall,and the high pressure turbine portion of the engine. For militaryaircraft, the high temperature reflector metal would also bebeneficially used in the augmentor portion of the engine.

[0012] An advantage of the present invention is that the radiation fromthe combustion process is reflected back into the gas flow path. Thisradiative heat, rather than being absorbed by the component in thecombustor or HPT portion of the engine, is absorbed by the fluid andcarried back into portions of the engine further aft that currentlyoperate at cooler temperatures. The result is that the component doesnot become as hot. At a given temperature of operation of the engine,the component, because it is operating at a cooler temperature, will notdeteriorate as rapidly due to thermal degradation, resulting in longercomponent life and less mean time between repair or refurbishment.

[0013] Another advantage of the present invention is that the fluidstream will be heated to a higher temperature as the reflected radiationis absorbed by the materials comprising the gaseous fluid and carriedfrom the combustor portion of the engine into the aft turbine portionsof the engine. This increased fluid temperature translates intoincreased engine efficiency, as the available energy in the fluid streamfor both extraction by the turbine to operate the engine and for thrustto propel the aircraft is greater.

[0014] Still another advantage of the present invention is that theengine can be operated at an even higher temperature than currentlyexperienced using the current invention if shortened component life andincreased repair rates can be tolerated in exchange for even greaterefficiency.

[0015] Other features and advantages of the present invention will beapparent from the following more detailed description of the preferredembodiment, taken in conjunction with the accompanying drawings whichillustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

[0016]FIG. 1 is a schematic representation of a high bypass turbofan gasturbine engine.

[0017]FIG. 2 is a schematic representation of a low bypass turbofan gasturbine engine equipped with an augmentor.

[0018]FIG. 3 is a schematic representation of the combustor and highpressure turbine sections of a gas turbine engine.

[0019]FIG. 4 is a cross-section of an as-manufactured high pressureturbine vane of a gas turbine engine after application of a conventionalthermal barrier system.

[0020]FIG. 5 is a cross-section of the high pressure turbine vane of thegas turbine engine after the outer surface of the ceramic topcoat hasbeen smoothed to achieve a surface finish of 50 micro inches or finer;and

[0021]FIG. 6 is schematic representation of the optical reflector of thepresent invention applied over a smooth ceramic topcoat.

DETAILED DESCRIPTION OF THE INVENTION

[0022] In accordance with the present invention, hot section componentsof a gas turbine engine which form the boundary of the gas flow path orwhich are located in the gas flow path are coated with a thin layer of aspecular optical reflective material that has a high temperaturecapability. The material as applied has a smooth surface finish so as toreflect the heat back into the fluid path and away from other hotsection components.

[0023] A high bypass aircraft gas turbine engine 10 is shownschematically in FIG. 1. During operation, air is forced through the fan12. A portion of the air bypasses the core of the engine and is used tocontribute to the thrust that propels the engine. A portion of the airis compressed in the booster 14 and compressor 16 portions of the engineup to 10-25 times atmospheric pressure, and adiabatically heated to800°-1250° F. in the process. This heated and compressed air is directedinto the combustor portion of the engine 18, where it is mixed with fuelsupplied through a fuel nozzle system 20. The fuel is ignited, and thecombustion process heats the gases to temperatures on the order of3200°-3400° F. These hot gases pass through the high pressure 22 and lowpressure 24 turbines, where rotating discs extract energy to drive thefan and compressor of the engine. The gases then are passed to theexhaust system 26, where they contribute to thrust for aircraftpropulsion.

[0024] Operation of a low bypass gas turbine engine, shown schematicallyat 30 in FIG. 2, is similar, except that operational requirements maydictate omission of the booster 14 and addition of an augmentor 28 inthe exhaust system shown at 26 in FIG. 1. To emphasize the conceptualsimilarity, the same identification numerals are employed in bothfigures.

[0025] The combustor 18 and high pressure turbine 22 sections of anengine such as in FIG. 1 or FIG. 2 are shown in greater detail in FIG.3. Compressed air from the compressor is introduced through a diffuser40 into an annular cavity defined by outer combustor case 42 and theinner combustor case 44. A portion of the compressed air passes througha swirl nozzle 46, where it is mixed with fuel supplied through a fueltube 48. The swirl nozzle and fuel tube are components of the fuelnozzle system 20. The fuel/air mixture is self-igniting under normaloperating conditions, except for those transient conditions where flameinstability or flame-out occurs. The flame is confined and directedtoward the turbine by the outer combustor liner 50 and the innercombustor liner 52. These liners are oriented about a central axis 55and are substantially symmetrical about this central axis 55 forming thegas flow path. Each combustor liner additionally is provided with aplurality of cooling holes 54, through which compressed air supplied bythe compressor is forced to pass. Compressed air in annulus betweenliners 50, 52 and combustor cases 42, 44 provides back side cooling toliners 50, 52 before exiting cooling holes 54. The combustor liners 50and 52 are described as having an inner side, and an outer side.

[0026] The hot gases of combustion then leave the combustor and enterthe high pressure turbine 22, which may comprise a single stage, asshown in FIG. 3, or multiple stages, each stage being comprised of anozzle 60 and a rotor 70. For the purposes of this discussion, thecombustor is presumed to be of a single stage configuration, forsimplicity of discussion, but the concepts of the present invention arefully applicable to gas turbines of other configurations and designswith additional turbine stages. The nozzle 60 is comprised of aplurality of vanes 62 disposed between and secured to an inner band 64and an outer band 66. Vanes 62 are substantially stationary, althoughthey may be capable of rotating about their axes in limited motion invariable guide vane configurations. The turbine nozzle 60 preferablyincludes a plurality of circumferentially adjoining segments 80collectively forming a complete 360° assembly. Each segment 80 has twoor more circumferentially spaced vanes 62 (one shown in FIG. 3), eachhaving an upstream leading edge and a downstream trailing edge, overwhich the combustion gases flow. As the temperature of the hot gas inthe gas flow path can easily exceed the melting point of the materialsforming the boundaries of the gas flow path, it is necessary to cool thecomponents forming the flow path, first by passing the air coming fromthe compressor at about 1000°-1250° F. (535°-675° C.) over the outersurfaces of the nozzles, then by using the same air after it passesthrough the cooling holes 102 (shown in FIGS. 4, 5, and 6) to direct athin film of air between the surface of the vanes 62 and the hot gases.The thin film of air forming a boundary layer assists in protecting thevanes 62 from being heated to even higher temperatures by a processreferred to as film cooling. Components of at least one turbine stageare often provided with cooling air through cooling holes. Additionally,the surface of the vanes 62 are also coated with thermal barrier coatingsystems, which are comprised of a bond coat applied between anunderlying superalloy base material and an overlying ceramic layer, tocreate a thermal barrier coating system that reduces the flow of heat tothe substrate material.

[0027] The rotor 70 is comprised of a plurality of blades, each havingan airfoil section 72 and a platform 74, which are securely attached tothe periphery of a rotating disk 78. Important associated structures tosupport the rotor are not shown. The blades cooperate with a stationaryshroud 76 to effect a gas seal between rotor 70 and the stationarycomponents of the engine.

[0028] Downstream of the fuel nozzle 46, the gas flow path is defined bythe inner surfaces of the inner combustor liner 52 and the outercombustor liner 50, and portions of the turbine or turbines includingthe inner and outer bands 64 and 66, the vanes 62, which direct the flowof gas, the airfoil 72, which extracts energy from the hot gas, theshrouds 76, as well as the exhaust system 26 and/or augmentor 28 aft ordownstream of the turbine section of the engine. The present inventionis specifically applicable to those components which define the gas flowpath downstream of the swirl nozzle 46. Systems for providing coolingair and thermal barrier coating systems are well-known in the gasturbine engine art.

[0029] Materials employed in the combustor, turbine and exhaust systemsections of aircraft gas turbines are typically high temperaturesuperalloys based on nickel, cobalt, iron or combinations thereof. Allof these superalloys are believed to be suitable substrate materials forthe present invention. Also, monolithic ceramic materials and fiberreinforced ceramic matrix composite materials, described hereincollectively as ceramic materials, may be employed in the combustor,turbine and exhaust systems sections of an aircraft gas turbine. Suchceramic materials are specifically contemplated for use in the presentinvention, and may have higher temperature limits than the hightemperature superalloys used for combustors.

[0030] Even for gas turbine engines designed for commercial airliners,gas velocity through the engine may approach the speed of sound. Thus,the total gas residence time in the engine is but a small fraction of asecond, during which time air coming through the compressor is mixedwith liquid fuel, and combustion of the mixture occur. As the mixture iscombusted to form a gas, heat, including radiant heat, is generated.Even with the most recent advances in cooling measures used in gasturbine engines such as active cooling controls and advanced thermalbarrier coating systems which reduce the amount and/or rate of heattransferred to components due to convective and conductive heattransfer, the temperatures of the components along the flow path surfaceare still elevated to very high temperatures. The present inventionassists in reducing the amount of heat transferred to these componentsby radiation transfer.

[0031] The present invention utilizes a high temperature specularoptical reflector applied directly over existing ceramic materials suchas thermal barrier systems utilized to protect the substrate material.These specular optical reflectors are applied as a very thin coating andin a manner so that they do not adversely affect the cooling holes inthe surfaces of the components along the gas flow path. Conventional andwell known techniques for applying thermal barrier coatings providesurfaces that are much too rough for the thin coatings to act as opticalreflectors. When these specular reflectors are applied over conventionalthermal barrier coatings having surface finishes of 100 micro inches andgreater, the rough surface causes the radiation to be scattered inmultiple of different directions and are ineffective in transferringheat back into the rapidly moving fluid. When the coatings are porous,such as when used for as a catalytic coating, the radiation isreabsorbed into the substrate, so it cannot be used as an opticalreflector.

[0032] In one embodiment of the present invention, the specular opticalcoating of the present invention is applied to the surface of anafterburner liner. In another embodiment of the present invention, thespecular optical coating of the present invention is applied to thesurface of a flameholder.

[0033] In another embodiment of the present invention, a turbine nozzleis manufactured in accordance with standard manufacturing methods. Astandard ray-tracing program could be used to optimize the geometry ofthe turbine nozzle to be coated with the specular optical coating of thepresent invention. In addition, there may be surfaces that have heatreflected onto them from reflection or refraction from neighboringturbine blades that could also benefit from the specular opticalcoating. Referring to FIG. 4, turbine vane 62 is comprised of asubstrate 110 suitable for use at high temperatures. As discussed above,the substrate can be selected from several materials. However, asillustrated in FIG. 4, substrate 110 is a high temperature nickel basesuperalloy. A bond coat 112 is applied over the nickel base alloysubstrate. Overlying bond coat 112 is a ceramic layer 114 having asurface 115 that has a rough surface finish. As used herein, the term“rough surface finish” is one that is greater than about 100 microinches. When the substrate is selected from one of the availabledifferent materials, such as a ceramic matrix composite material, thebond coat 112 may be omitted.

[0034] The surface finish of the thermal barrier coating system istypically too rough to act as a specular optical reflector because ofthe manufacturing techniques used to apply the ceramic top coat. Giventhe complex geometry of the turbine nozzle and vane, the forward facingsurface of the upstream leading edge of the turbine vane, that is, theouter surface of the thermal barrier coating overlying the substratesurface, is then polished. In one embodiment, the forward facing surfaceof the upstream leading edge of the turbine vane is polished by handusing fine emery paper so that the surface 115 of the ceramic layer 114,as shown in FIG. 5, has a surface finish of no greater than about 50micro inches, preferably about 32 micro inches and smoother. It isundesirable to polish the other surfaces of the turbine nozzle or tocoat the other surfaces of the turbine nozzle with the coating of thepresent invention as such polishing and coating would only serve toreflect the heat of the gas flow path into a nearby vane or wall. Thissmooth surface of the forward facing surface of the turbine nozzleleading edge is required to achieve the reflective properties requiredfor the present invention to be effective. Additionally, the smoothsurface assists in maintaining a smooth laminar-like flow of the coolinglayer adjacent to the surface of the component by minimizing turbulence.In production, well known polishing techniques such as lapper wheelswith diamond paste and tumbling can be employed to speed the polishingprocess and increase throughput.

[0035] Next, the forward facing surface of the upstream leading edge ofthe turbine nozzle vane is coated with a very thin specular reflectivecoating 116 of a material, as shown in FIG. 6, that will reflect theradiation away from the surface. A standard ray-tracing program could beused to optimize the geometry of the turbine nozzle vane to be coatedwith the specular optical coating of the present invention. In addition,certain surfaces of the turbine nozzle vanes have heat reflected ontothem from reflection or refraction from neighboring turbine blades thatcould also benefit from the specular optical coating. The coating 116 isapplied by a process that deposits material so that a very smoothsurface finish is maintained. A preferred method is a chemical vapordeposition (CVD) process that deposits a coating to a thickness of about1 micron (0.0004″). Other acceptable methods for depositing this thinspecular coating to a thickness of about 1 micron include sputtering,liquid phase infiltration and physical vapor deposition. However, notall methods for depositing a coating produce coatings consistent withthis invention. Other methods such as thermal spray methods do notproduce an acceptable coating for specular reflection, as the coatingsdeposited by these processes are too thick and too rough. The thicknessof the specular layer can be greater, for example 10 microns or less,but is maintained at about 1 micron because of the great expense of thematerial used as the specular reflector.

[0036] A preferred specular reflector coating material is platinum,although palladium or multiple dielectric mirrors comprising tantalumoxide (Ta₂O₅), silica (SiO₂), titanium dioxide (TiO₂) and combinationsthereof. It is fundamental that the material used as a coating materialremain highly reflective as the hot gas stream 120 passes over thesurface. Thus, thick non-adherent oxide scales cannot form, as theformation of these scales destroy the effectiveness of the coating as areflector. Also, the very thin coating, in addition to being lessexpensive, is extremely adherent to the polished TBC, and, due to itsthinness, does not peel off in layers, which peeling can adverselyaffect the surface finish. The thin layer does not provide a severeweight penalty for the components to which it is added. In addition, thelayer is maintained as a thin layer to allow the surface finish to be ofhigh reflective, optical quality.

[0037] Testing of other reflective combustor components has indicatedthat a specular reflective layer can reflect at least about 80% of theincident radiation, an amount of radiation sufficient to lower thetemperature of a component by up to about 100° F. when the temperatureof a ceramic coating adjacent to the fluid stream is at 2300° F. (1260°C.) as compared to a component having a ceramic coating but without thespecular reflective layer. These components have displayed animprovement of 95° F., as measured by thermocouples attached todeflectors in a high pressure sector test for approximately 100 hours,as compared to a substantially identical deflector that lacked a coatingsuch as described by the present invention.

[0038] While the present invention has been described as an improvementto a turbine nozzle, the present invention can be applied to any othersurface along the gas flow path of a turbine engine or other hightemperature devices, such as a continuous furnace or a burner. Forexample, the specular reflective coating can be applied to the combustorwalls, so that any incident radiation is reflected away from thecombustor walls and into the gas flow path. Because at least a portionof the energy is reflected from the components comprising the gas flowpath, thereby lowering their temperature, the radiation is absorbed bythe gases in the gas flow path, thereby raising its temperature.

[0039] While the invention has been described with reference to apreferred embodiment, it will be understood by those skilled in the artthat various changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment disclosed as the best modecontemplated for carrying out this invention, but that the inventionwill include all embodiments falling within the scope of the appendedclaims.

1. A component having a specular reflective surface for use in a hotflow path of a high temperature device in which hot gaseous fluidstraverse the device in a hot flow path, the component comprising: asubstrate material having a surface that forms a boundary for hotgaseous fluids of combustion; and a specular reflective coating having apredetermined thickness overlying the substrate surface forming the hotfluid boundary, the specular reflective material having an appliedroughness sufficiently smooth, and having a high temperature capabilityto survive temperatures in the hot flow path so that a surface of thespecular reflective material reflects a least about 80% of incidentradiation away from its surface to the fluids in the hot flow path andwherein the coating is not subject to oxidation due to contact with thehot gaseous fluids.
 2. The component of claim 1 wherein the device is agas turbine engine.
 3. The component of claim 1 wherein the component isan afterburner liner.
 4. The component of claim 1 wherein the componentis a flameholder.
 5. The component of claim 1 further including aceramic material between the substrate material and the specularreflective coating, the ceramic material forming a thermal barrieroverlying the substrate material, a surface of the ceramic materialopposite the substrate and adjacent the specular reflective coatinghaving a surface roughness about 50 micro inches and smoother.
 6. Aturbine nozzle component having a specular reflective surface for use ina hot flow path of a high temperature device in which hot gaseous fluidstraverse the device in a hot flow path, the component comprising: asubstrate material wherein the specular reflective surface forms aboundary for hot gaseous fluids of combustion; and a specular reflectivecoating having a predetermined thickness overlying the substrate surfaceforming the hot fluid boundary, the specular reflective material havingan applied roughness sufficiently smooth, and having a high temperaturecapability so that a surface of the specular reflective materialreflects a least about 80% of incident radiation away from its surfaceto the gases in the hot flow path and wherein the coating is not subjectto oxidation due to contact with the hot gaseous fluids.
 7. The turbinenozzle component of claim 6 wherein the component is a turbine vane. 8.The turbine nozzle component of claim 6 further including a ceramicmaterial between the substrate material and the specular reflectivecoating, the ceramic material forming a thermal barrier overlying thesubstrate material, a surface of the ceramic material opposite thesubstrate and adjacent the specular reflective coating having a surfaceroughness about 100 micro inches and smoother and the surface of thespecular reflective coating having a surface finish of 100 micro inchesand smoother.
 9. The turbine nozzle component of claim 8 wherein theturbine nozzle component is a turbine nozzle vane wherein the substratematerial has a leading edge and a trailing edge and wherein the specularreflective coating is only applied to the leading edge.
 10. The turbinenozzle component of claim 9 wherein the surface of the ceramic materialopposite the substrate and adjacent the specular reflective coating hasa surface roughness about 50 micro inches and smoother.
 11. The turbinenozzle component of claim 9 wherein the surface of the ceramic materialopposite the substrate and adjacent the specular reflective coating hasa surface roughness about 32 micro inches and smoother.
 12. The turbinenozzle component of claim 9 wherein the specular reflective coating isselected from the group of materials consisting of platinum, palladium,a dielectric mirror comprising tantalum oxide (TaO₂), a dielectricmirror comprising silica (SiO₂), a dielectric mirror comprising titaniumdioxide (TiO₂) and combinations thereof.
 13. The turbine nozzlecomponent of claim 9 wherein the coating is applied to a predeterminedthickness up to about 10 microns.
 14. The turbine component of claim 9wherein the coating is applied to a predetermined thickness about 1micron and less, and forms a continuous coating.
 15. The component ofclaim 14 further characterized by a temperature performance improvementof about 100° F. over a component without the specular coating.
 16. Thecomponent of claim 6 wherein the substrate material is a hightemperature superalloy selected from the group consisting of asuperalloy selected from the group consisting of Fe, Co, Ni andcombinations thereof.
 17. A method for manufacturing a turbine nozzlehaving a specular reflective surface for use in a hot flow path of a gasturbine engine, the method comprising the steps of: providing a turbinenozzle vane having a leading edge and a trailing edge comprised of asubstrate material; applying a ceramic thermal barrier coating systemover the substrate surface of at least the leading edge of the turbinenozzle vane forming the hot fluid boundary; mechanically working asurface of a ceramic coating forming the outer layer of the thermalbarrier coating system, overlying and opposite the substrate surface toobtain a surface finish of about 100 micro inches and smoother; applyinga specular reflective coating over the surface of the ceramic coating toa predetermined thickness, the method for applying the coating providinga coating surface finish of about 100 micro inches and smoother, anouter surface of the specular reflective coating opposite the ceramiccoating being exposed to gaseous fluids in the hot flow path of theengine.
 18. The method of claim 17 wherein the step of mechanicallyworking the surface of the ceramic coating further includes obtaining asurface finish of about 50 micro inches and smoother.
 19. The method ofclaim 17 wherein the step of mechanically working the surface of theceramic coating further includes obtaining a surface finish of about 32micro inches and smoother.
 20. The method of claim 17 wherein the stepof applying the specular reflective coating over the surface of theceramic coating to a predetermined thickness includes applying thecoating to a thickness of up to about 10 microns.